Turbine propulsion-gas generator for aircraft and the like



1968 c. c. E. DECHAUX TURBINE PROPULSION-GAS GENERATOR FOR AIRCRAFT AND THE LIKE 4 Sheets-Sheet 1 Filed May 9, 1966 NM t.

3 muA /N VE NT OR. Charles C. E. Dech crux Attorney Aug'. 20, 1968 c. c. E. DECHAUX 3,397,535

TURBINE PROPULSION-GAS GENERATOR FOR AIRCRAFT AND THE LIKE 4 Sheets-Sheet 2 Filed May 9. 1966 INVENT OR. Charles C. E. Dec'houx BY Attorney 3,397,535 CRAFT 1968 c. c. E. DECHAUX TURBINE PROPULSIONGAS GENERATOR FOR AIR AND THE LIKE 4 Sheets-Sheet 5 Filed May 9, 1966 F795 INVENTOR.

Charles C. E. Dechaux BY Attorney Aug. 20, 1968 c. c. E. DECHAUX 3,397,535

TURBINE PROPULSION-GAS GENERATOR FOR AIRCRAFT AND THE LIKE 4 Sheets-Sheet 4 Filed May 9. 1966 INVENTOR. Charles C.E.Dechaux BY A ttorn e y United States Patent 3,397,535 TURBINE PROPULSION-GAS GENERATOR FOR AIRCRAFT AND THE LIKE Charles Camille Emile Dechaux, 11 Rue Francis de Pressence, Chatenay-Malabry, France Filed May 9, 1966, Ser. No. 548,590 7 Claims. (Cl. 60-395) ABSTRACT OF THE DISCLOSURE A turbine propulsion-gas generator having an elongated housing enclosing a primary compressor whose flow of primary air is fed to a combustion chamber receiving fuel from an injector, the resulting hot combustion gas being directed at a turbine prior to emergence at an outlet at the periphery of said housing ahead of said turbine, while a secondary compressor induces a flow of secondary air into the housing and outwardly through the outlet, a rotatable partition being disposed in the housing between the turbine and the secondary compressor for limiting interaction between the combustion gas and the secondary air flowing to the outlet means to a peripheral region of the housing.

This invention relates to turboprops or internal-combustion engines comp-rising compressors and turbines adapted to deliver a considerable volume of gas under a relatively reduced pressure.

A turboprop or turbopropulsion engine of this character is particularly useful for aircraft applications, notably for driving the main rotors of helicopters or other similar rotorcraft.

The turboprop according to this invention, comprising in a common body at least one combustion chamber, a fuel distributor, ignition means, a primary air compressor feeding said chamber, a turbine inserted in the exhaust circuit of said chamber, a secondary air compressor, said primary and secondary compressors and said turbine being operated by a common longitudinal axial shaft, and an exhaust duct, is characterized in that a partition is disposed across said body between the turbine and the secondary air compressor, said partition being notched along its outer periphery so that the gas stream and the secondary air stream can contact each other only in said peripheral zone.

Other features and advantages of the invention will appear as the following description proceeds with reference to the accompanying drawing given by way of example in order to aiford a clearer understanding of this invention and of the manner in which the same may be carried out in practice. In the drawing:

FIG. 1 is a longitudinal axial section of the turboprop gas generator constructed according to the teachings of this invention;

FIG. 2 is a section taken upon the line IIII of FIG.

FIG. 3 is a section taken upon the line III-III of FIG. 1;

FIG. 4 is a section taken upon the line IVIV of FIG. 1;

FIG. 5 is a section taken upon the line V-V of FIG. 1;

FIG. 6 is a section taken upon the line VI-VI of FIG. 1; and

FIG. 7 is an end view.

As shown in the figures, the turboprop according to this invention comprises a sheet-metal tubular structure 1 in the form of a body of revolution consisting of two bulbiform end portions 1a and 1c interconnected by an intermediate waist-like section 1b.

This body advantageously consists of three juxtaposed elements assembled together by means of assembly flanges such as If coincident with the major diametral plane of each bulb 1a and 1c. Sandwiched between the flanges of the sections constituting the bulb 1a is the peripheral edge of a transverse partition 10 covering the whole of the central portion and supporting a peripheral set of airfoil blades, FIGS. 1 and 3, slightly inclined to the corresponding radial planes and extending towards the interior of the intermediate waist-like section 1b. Owing to their inclinations these blades partly overlap one another and provide therebetween passages permitting the communication between the chambers separated by said partition 10.

This partition 10 actually consists of two walls providing therebetween an auxiliary tank R communicating through pipe lines (not shown) with the main fuel tank. One of these walls carries nozzle means 12 for ejecting the fuel under pressure.

Arranged in the outer portion of bulb 1a is a set of fixed inclined airfoil blades 14 serving a purpose to be explained presently.

Mounted in the waist-like central section 1b of the tubular housing is an annular combustion chamber Cb. This chamber is bounded by two surfaces of revolution consisting of circular partitions 15 and 16 forming an inner or central passage Ac and an outer or peripheral annular passage Ap, both passages or channels Ac and Ap permitting the flow of cooling air along the walls of chamber Cb. Between the partitions 15 and 16 an annular inlet Ba and an outlet divided into a plurality of nozzles 8 (FIGURE 4) are formed, these nozzles 8 being separated from one another by curved radial partitions 17 providing air passages At (FIG. 4) therebetween.

The bulb 10 carries the tangential gas-exhaust duct 18 and the body 1 comprises at each end an axial air-intake aperture Ed and Eg having each a spider-like support 19, 19' mounted therein for supporting a corresponding rotary-contact bearing 20, 20'.

These bearings 20, 20' have journaled therein the ends of a longitudinal shaft 2 having keyed thereon the hubs 3, 5, spaced by a sleeve 4' carrying the further hub 4.

The first hub 3 carries on the one hand an integral flange 6 rotating in a plane parallel to the aforesaid fixed partition 10, and a first series of relatively large radial airfoil blades 7 (FIGS. 1 and 2) extending at right angles to said flange 6 from the center to within a short distance from the inner edges of the fixed blades 14, and on the other hand another series of helical airfoil blades 28 mounted on the outer peripheral edge of flange 6 between adjacent radial blades 7.

The outermost section of the bulbiform end portion 1a, including the fixed blades 14 and the rotor revolving therein which consists of the hub 3 and the blades 7 and 28, constitutes the primary-air compressor delivering a relatively small volume of air at a relatively high pres-v sure, this air constituting the combustion medium supplied to chamber Cb and at the same time the air stream for cooling the walls of this chamber. This compressor is of the two-stage type, the first stage consisting of blades 7 and the second one of blades 28. Of course, a greater number of stages may be contemplated if desired.

The intermediate hub 4 is connected through a set of helical blades 4a to an annular element 4b, the outer ends of these blades registering with the outlets of nozzles 12 and the inlet of chamber Cb. This annular element constitutes an atomizer adapted to divide the fuel jet issuing from the nozzles and to mix same with the primary combustion air. The blades 4a act somewhat in the fashion of a fan to force the air through the slightly tapered passage Ac.

The opposite or downstream end hub 5 carries an integral plate 9 rotating in a plane at right angles to the axis of shaft 2; this plate carries on each face a set of blades, i.e., on the face adjacent the combustion chamber a first central set of radial blades 21 (FIGS. 1 and 5) for the purpose of forcing cooling air towards the curved partitions 17 (FIG. 4) separating the nozzles 8 from one another, and a second set of inclined blades 22 (FIGS. 1 and 5) registering with said nozzles 8 on the upstream side of disk 9.

These two sets of blades constitute the power turbine upon which the nozzles 8 are trained. On the opposite face of disk 9 (FIGS. 1 and 6) the hub 5 carries a first central set of radial blades 23 and a second set of peripheral inclined blades 24.

The blades 23 constitute the low-pressure compressor rotor delivering a relatively large volume of air. This compressor may also be of the single-stage or multi-stage type, as desired.

The blades 23 and 24 are advantageously constituted by a continuous plane surface and, as Will be seen from FIGS. 1 and 5, form a set of integral vanes With the up stream blades 21, 22, these vanes reaching around the periphery of disk 9 and subdividing the annular clearance between the disk and the tubular housing wall into a plurality of peripherally juxtaposed, mobile compartments C (FIG. 6).

It will be noted that the outlet (or rather both outlets 18) is disposed tangentially to the bulbiform end portion so as to substantially overlap the fluid stream delivered by both sets of blades 23 and 24. Thus, the vane structure 21-24 sweeps the inner wall surface of the engine housing 1 on both sides of the partitioning disk 9, over a region which includes the outlets 18, so as to form guide paths for positively channeling the combustion gas from nozzles 8 and the secondary air from intake 19' together with the cooling air exiting from channels Ap and Cb-into the compartments C where these fluids intermingle before being discharged from the housing.

The turboprop described hereinabove operates as follows:

The shaft 2 driven from turbine 22 rigid with hub 5 drives in turn the high-pressure compressor 3, 7, 28 delivering combustion air and cooling air. The nozzles 8 are cooled by the low-pressure compressor constituted by the blades 21. The thrust thus created is applied to the turbine blades 22. To obtain the maximum value of this gaseous thrust the direct passage towards the outlet 18 (which would constitute a leakage) is counteracted by the flow delivered by the low-pressure compressor 23 which constitutes an air cushion impinging against the thrust gas in the zone of the integral blade structure 22-24. The formation of this gaseous pad between the blades, i.e., within the peripheral compartments C defined thereby, is particularly important because it controls the actual elficiency of the turbine.

In turboprop machines of known type (such as the well-known Oryx turboprop engine constructed by Napier) the gas stream issuing from the turbine also impinges against the fluid stream forced by the auxiliary compressor towards the outlet zone; however, the efficiency of this apparatus is disputable for the gaseous streams impingev against each other throughout the crosssectional area of the engine, and act in opposition to each other.

As contrasted thereto, in the case of the turboprop according to this invention the combined thrusts of the combustion gas and of the secondary air are applied in the same direction against the common blades 22-24 without any possibility of interpenetrating each other in the central portion.

On the other hand, the hot gases draw the cold air from the compressor as they meet, thus increasing the compressor speed and reducing its power demand.

Of course, it would not constitute a departure from the subject matter of the present invention to devise any modifications and variations of the specific embodiment illustrated, described and suggested herein, notably by applying the turboprop to the propulsion of machines such as helicopters, aircraft, hovercraft and the like.

What I claim is:

1. A pressure-gas-generating turbopropulsion engine comprising an elongated housing with a tubular wall centered on an axis and with first and second air-intake means at opposite ends thereof; a primary compressor in said housing adjacent said first air-intake means; a combustion chamber formed in said housing downstream of said primar-y compressor for receiving therefrom a stream of primary air aspirated through said first air-intake means; fuel-injection means in said housing for delivering fuel to said combustion chamber in admixture with said stream of primary air whereby a hot combustion gas is produced in said chamber upon ignition of the air/ fuel mixture therein; a turbine rotatably coupled With said primary compressor and mounted in said housing downstream of said combustion chamber for rotation about said axis by the combustion gas emerging from said combustion chamber; outlet means at said tubular wall downstream of said turbine for discharging the combustion gas from said housing in the wake of the rotating turbine; a secondary compressor downstream of said turbine adjacent said second air-intake means and rotatably coupled with said turbine to induce a flow of secondary air into said housing; transverse partition means rotatably coupled with and interposed between said turbine and said secondary compressor, said partition means being centered on said axis and having a periphery spaced from said tubular wall to define therewith an annular clearance communicating with said outlet means; and a set of generally radial vanes mounted on the upstream and downstream sides of said partition means and extending around the latter through said clearance, and vanes subdividing said clearance into a plurality of peripherally juxtaposed mobile compartments and forming guide paths for channeling said secondary air and said combustion gas into said compartments for intermingling therein prior to dis charge through said outlet means.

2. An engine as defined in claim 1 wherein said turbine and said compressors are provided with a common shaft extending axially through said housing, said partition means being a disk rigid with said shaft, said secondary compressor comprising a set of blades on said shaft integral with said vanes on the downstream side of said disk.

3. An engine as defined in claim 2 wherein said vanes extend from said shaft to the inner wall surface of said housing over a region extending both upstream and downstream of said disk, said outlet means being disposed in said region.

4. An engine as defined in claim 3 wherein said housing is formed with a restricted waist portion around said combustion chamber and with an enlarged end portion around said disk, said outlet means being disposed on said end portion.

5. An engine as defined in claim 4 wherein said combustion chamber is spaced from said tubular wall and defines therewith an annular channel for the direct passage of primary air from said first air-intake means through said waist portion and said compartments to said outlet means.

6. An engine as defined in claim 3 wherein said vanes terminate on said upstream side in a set of second-stage compressor blades close to said shaft, said combustion chamber forming with said shaft an annular central channel for the direct passage of primary air from said first air-intake means past said second-stage compressor blades through said compartments to said outlet means.

7. An engine as defined in claim 6 wherein said combustion chamber is spaced from said tubular Wall and defines therewith an annular peripheral channel for the direct passage of primary air through said compartments 6 to said outlet means, said combustion chamber terminating in a set of nozzles trained toward said compartments between the downstream ends of said central and peripheral channels.

References Cited UNITED STATES PATENTS 1,595,278 8/1926 Woerner 60-39.5 X 2,625,794 1/1953 Williams et al 6039.65 2,635,804 4/1953 Jedrzykowski 230116 2,891,382 6/1959 Brofiitt 60-39.66

JULIUS E. WEST, Primary Examiner. 

